Attitude control of a helicopter, e.g. banking, most often occurs by cyclic pitch control where the pitch angle of the rotary blade is continuously varied during the movement of the blade in the plane of rotation. However, this is mechanically a relatively complicated system and takes up substantial weight, which is a disadvantage, especially, for small helicopters. In addition, to control the yaw moment, a tail rotor is, normally, used, making the system even more complex. Therefore, for light-weight helicopters, different systems have been proposed.
In International patent application WO 99/38769, an unmanned helicopter is disclosed, where cyclic pitch control has been omitted as well as a tail rotor. The fuselage is connected to the rotor via a cardan-like joint for permitting limited tilting of the rotor relative to the fuselage in the roll and pitch direction.
In U.S. Pat. No. 7,128,293, a different pivot coupling between a cabin and a power unit with a rotor. The rotor can be moved forward and rearward by a trim actuator to trim the helicopter and it is tilted about pitch and roll axes by directional actuators for directional control.
In British patent application GB2375090 a three-leg suspension for the rotor is disclosed.
When a rotor is tilted relatively to a fuselage, the gyroscopic forces make control of pitch and roll of the helicopter difficult, especially, if the rotor is without cyclic pitch. For this reason, it has been proposed to use two counter-rotating rotors.
In U.S. Pat. No. 6,460,802, a helicopter with a double-rotor is disclosed. The double rotor is tiltable for controlling pitch and roll. A tiltable double rotor is also disclosed in International patent application WO 04/130814. International patent application WO 04/002824 discloses a helicopter with two tiltable rotors, one rotor provided on each end of the aircraft.
Especially for helicopters without cyclic pitch, any tilting of the rotor, especially if not a counter-rotating double-rotor, induces critical forces causing unstable pitch and roll of the aircraft. Precession induced by a tilting of the rotor is naturally damped due to the gravity forces on the aircraft. However, for large and fast tilting of the rotor, the resulting stronger gyroscopic forces on the aircraft with a corresponding precession may be fatal. It would be desirable to find an improved, simple, and light-weight solution in this respect.